CN1459550A - Method and device for enhancing use life of gas turbine wing surface - Google Patents

Method and device for enhancing use life of gas turbine wing surface Download PDF

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Publication number
CN1459550A
CN1459550A CN03136888A CN03136888A CN1459550A CN 1459550 A CN1459550 A CN 1459550A CN 03136888 A CN03136888 A CN 03136888A CN 03136888 A CN03136888 A CN 03136888A CN 1459550 A CN1459550 A CN 1459550A
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China
Prior art keywords
blade
tenon
shank
cooling cavity
width
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Granted
Application number
CN03136888A
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Chinese (zh)
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CN100572757C (en
Inventor
J·P·海瓦德
C·A·弗莱克三世
T·L·诺里斯
T·S·赫夫伦
R·D·乌斯特曼
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General Electric Co
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General Electric Co
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/085—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor
    • F01D5/087—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor in the radial passages of the rotor disc
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2240/00—Components
    • F05B2240/80—Platforms for stationary or moving blades
    • F05B2240/801—Platforms for stationary or moving blades cooled platforms
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms

Abstract

A method enables a gas turbine engine blade to be manufactured to include an airfoil, a platform, a shank, and a dovetail. The platform extends between the airfoil and the shank, the shank extends between the dovetail and the platform, and the dovetail includes at least one tang for securing the blade within the engine. The method comprises defining a cooling cavity in the blade that extends through the airfoil, the platform, the shank, and the dovetail, such that a portion of the cavity defined within the dovetail includes a root passage portion having a first width, and a transition portion extending between the root passage and the portion of the cavity defined within the shank, and wherein the portion of the cavity defined within the shank has a second width that is larger than the root passage first width.

Description

Improve the gas turbine airfoil method and apparatus in working life
Background of invention
The present invention relates generally to gas turbine, more specifically, relates to the turbine blade that uses in gas turbine.
At least some known gas turbines have a core-engine, this motor is furnished with the high-pressure compressor that is used to compress the air-flow that enters motor, the firing chamber that is used for combustion fuel and AIR MIXTURES and turbine successively by flowing of fluid, this turbine has a plurality of rotor blades, and they draw rotational energy from the ignition mixture air-flow of being discharged by the firing chamber.Because turbine will bear the high temperature gas flow of discharging from the firing chamber, so the parts of turbine will cool off the thermal stress that is produced by high temperature gas flow with the reduction meeting.
Rotation blade has the aerofoil of hollow, and it is used for providing cooling air by cooling circuit.Aerofoil comprises the cooling cavity that is defined by sidewall, and sidewall defines this cooling cavity.The cooling of the cooling raio of engine components such as the parts of high-pressure turbine is essential, because employed material has the thermal stress restriction in the structure of such parts.Generally, draw cooling air from the outlet of gas compressor, and cooling air is used to cool off for example turbine airfoil.Cooling air enters blast tube downstream, firing chamber again after the cooling turbine aerofoil.
At least some known turbine airfoils have the cooling circuit that is used to guide the cooling air flow that cools off aerofoil.More specifically, the cavity in the aerofoil defines the runner that is used to guide cooling air.Such cavity can limit, and for example, has the serpentine channel of multiple exit.Cooling air enters serpentine channel via the root of aerofoil.In the design of at least some known aerofoils, the transition that a sudden change arranged between root and airfoil portion to be increasing the cross sectional area of cooling cavity, with the increase of the capacity of the cooling air that enters airfoil portion.Because the thermal stress internal cavities that can lead, so coating environment coating is beneficial to prevent to cool off oxidation in the cavity on the wall that limits cavity.Because the restriction of the geometrical shape of cooling channel, in coating process, coating also can be deposited in the root of aerofoil.
In order to help resisting built-in thermal stress, at least some known blades apply a layer thickness and approximately equal 0.001 inch environment coating.Environment coating with thickness like this prevents from the oxidation of cavity wall and helps aerofoil to keep out thermal stress and the mechanical stress that produces in the higher running temperature zone of blade.Yet if coated coating is thicker, the acting in conjunction meeting of the suddenly transition in the thickness of the increase of environment coating and the tenon makes the root at aerofoil occur crackle prematurely, the transition zone of tenon because stress can lead.Continuously operation will cause the too early ring of ruining of motor intra vane after a period of time.
The invention summary
An aspect of of the present present invention provides a kind of manufacture method of blade of gas turbine.This blade has an aerofoil, a platform, a shank and a tenon, and wherein platform extends between aerofoil and shank, and shank extends between tenon and platform, and tenon comprises that at least one is used for vanes fixed at in-engine tang.This method is included in and limits a cooling cavity in the blade, it extends through aerofoil, platform, shank and tenon, wherein be limited to cavity in the tenon and partly have the root channel part of first width, and one at the root passage be limited to the transition portion that extends between the cavity part in the shank, is limited to second width that the cavity part in the shank had first width greater than the root passage.This method also is included at least a portion of internal surface of the blade that limits this cooling cavity and applies the anti-oxidant environment coating of one deck.
Another aspect of the present invention provides a kind of gas turbine blades.This blade has a platform, a shank that extends from this platform, and a tenon that between end of blade and shank, extends, is used in gas turbine, installing blade, and wherein this tenon has at least one tang.This blade also has an aerofoil, this aerofoil has a first side wall and one second sidewall that extends between platform and blade tip in radial extension, with one by tenon, shank, platform and aerofoil are limited to the cooling cavity in the blade, this cooling cavity has a tenon part that is limited in the tenon, one is limited to a shank part and the airfoil portion that is limited in the aerofoil in shank and the platform, wherein shank part flows with their between airfoil portion and tenon part and links to each other, tenon partly comprises a passage and a transition passage, the root passage has one first width, shank partly has second width greater than above-mentioned first width, and transition passage is connected between root passage and the shank part.
One side more of the present invention provides a kind of gas turbine with a plurality of blades.Each blade has an aerofoil, a shank and a platform that extends between aerofoil and shank.Each blade also has a cooling cavity and and has at least one and be used for the tenon of blade at the tang of motor internal fixation.Shank extends between platform and tenon, and the cooling cavity is limited by aerofoil, platform, shank and tenon, and have tenon part, shank part with its airfoil portion that flows and link to each other.Tenon partly has a root passage and the transition passage that its width is first width.Shank partly has one second width, and it is greater than first width of root passage, and transition passage is tapered between root passage and shank part.
The accompanying drawing summary
Fig. 1 schematically shows a gas turbine;
Fig. 2 is the perspective view that is used for the turbine rotor device of gas turbine shown in Figure 1;
Fig. 3 is the sectional side view that is used for the rotor blade of rotor arrangement shown in Figure 2;
Fig. 4 is the sectional front view of rotor blade shown in Figure 3; With
Fig. 5 is the sectional front view of the part of prior art rotor blade.
Detailed description of the Invention
The gas turbine 10 that Fig. 1 schematically illustrates comprises a blower fan apparatus 12, high-pressure compressor 14 and a firing chamber 16.Motor 10 also comprises a high-pressure turbine 18 and a low-pressure turbine 20.Motor 10 has an inlet side 28 and an outlet side 30.In one embodiment, motor 10 is the CFM-56 h type engine hs that can buy from the CFM international corporation of Ohioan Cincinnati.
In the actual motion, air stream is by blower fan apparatus 12, and air is infeeded in the high-pressure compressor 14 and is compressed.Be admitted to firing chamber 16 by the air of high compression.From the air stream drives turbine 18 and 20 of firing chamber 16, and turbine 20 drives blower fan apparatus 12.Turbine 18 drives high-pressure compressor 14.
Rotor arrangement 40 shown in the perspective view of Fig. 2 can be used for gas turbine, such as gas turbine 10 (as shown in Figure 1).Device 40 comprises a plurality of rotor blades 42 that are installed in the rotor disk 44.In one embodiment, blade 42 forms a High Pressure Turbine Rotor leaf-level (not shown) of gas turbine 10.
Rotor blade 42 extends radially outward from rotor disk 44, and each comprises an aerofoil 50, a platform 52, a shank 54 and a tenon 56.Each aerofoil 50 comprises the first side wall 60 and second sidewall 62.The first side wall 60 is convexs, defines the suction side of aerofoil 50, and second sidewall 62 is spills, defines aerofoil 50 on the pressure side.Sidewall 60 and 62 couples together at the guide lug 64 of aerofoil 50 and trailing edge 66 places axially spaced with it.More particularly, aerofoil trailing edge 66 is positioned at downstream and its position at a distance of chord length of aerofoil guide lug 64.
First and second sidewalls 60 and 62 respectively from the blade root 68 that closes on platform 52 to outer length or radially extend to aerofoil top 70.Aerofoil top 70 defines the radially external boundary (not shown in Fig. 2) of an inner cooling chamber.Cooling chamber is defined in the aerofoil 50, between the sidewall 60 and 62, and extends through platform 52 and shank 54 enters tenon 56.More specifically, aerofoil 50 comprises an internal surface (not shown among Fig. 2) and an outer surface 74, and cooling chamber is limited by the aerofoil internal surface.
Platform 52 extends between aerofoil 50 and shank 54, makes each aerofoil 50 extend radially outwardly from each platform 52.Shank 54 extends radially inwardly to tenon 56 from platform 52.Tenon 56 extends radially inwardly from shank 54 and is beneficial to rotor blade 42 is fixed on the rotor disk 44.More specifically, each tenon 56 comprises at least one tang 80, and this tang 80 extends radially outwardly from tenon 56, is beneficial to install in each tongue-and-groove 82 each tenon 56.In this embodiment, tenon 56 comprise a upper blade tang to 84 and one bottom blade tang to 86.
Fig. 3 is the sectional view of the local guide lug of rotor blade 42.Fig. 4 is the side partial cross-sectional of rotor blade 42.Fig. 5 is the sectional view of the part of prior art rotor blade 100.Each blade 42 comprises platform 52, shank 54 and tenon 56.As mentioned above, shank 54 extends between platform 52 and tenon 56, and tenon 56 extends radially inwardly to a radial inner end 101 of blade 42 from shank 54.Platform 52, shank 54, tenon 56 and aerofoil 50 are hollow, and define a cooling cavity 102 that connects.More specifically, cooling cavity 102 is defined in the rotor blade 42 by the internal surface of blade 42.Cooling cavity 102 comprises a plurality of inwalls 106, and they will cool off cavity 102 and be separated into a plurality of cooling chambers 108.Internal connection between chamber 108 and the wall 106 and geometrical shape change according to the application target of blade 42.In one embodiment, interior wall 106 and aerofoil 50 cast inblocks.
Blade cooling cavity 102 also comprises a tenon part 112, a shank part 114 and an airfoil portion 116 that flows and link to each other, and the cooling fluid of sending into cooling cavity tenon part 112 like this enters cooling cavity airfoil portion 116 by part 112 and 114.Cooling cavity tenon part 112 comprises a blade root channel section 120 and a transition passage section 122 that flows and link to each other.More specifically, blade root channel section 120 has a plurality of root passages 124 that extend between blade end 101 and transition passage section 122, and transition passage section 122 is extended between blade root channel section 120 and shank part 114.
The width D of blade root channel section 120 RConstant substantially, this width is at the suction side 132 of cooling cavity 102 and on the pressure side measures between 134.More specifically, the width D of between the radial outer end 140 of the radial inner end 138 of blade root channel section 120 and blade root channel section 120, measuring ROn length 136, be substantially invariable.Blade root channel section radial inner end 138 is near cooling cavity throat 141, and blade root channel section radial outer end 140 is near transition passage section 122.Cooling cavity throat 141 blade end 101 be limited to the lower blade tang to 86 between, blade root channel section radial outer end 140 is limited to the upper blade tang between 84.Therefore, sidewall 132 and 134 is parallel basically in blade root channel section 120.
Transition passage section 122 is outwards tapered gradually to cooling off cavity shank part 114, the width D of cooling cavity shank part 114 from blade root channel section 120 SGreater than blade root channel section width D RTherefore, the width D of transition passage section 122 TBetween the radial outer end 144 of transition passage section 122 and radial inner end 142, change.The transition passage section width D that changes TIn whole transition passage section 122 greater than blade root channel section width D R, and equal shank partial width D at transition passage radial outer end 144 places SThe length 146 of transition passage section 122 measures between the end 142 and 144 of transition passage section.More particularly, transition passage segment length 146 with have predetermined radii and be limited at transition passage section 122 and blade root channel section 120 between the combining of arc-shaped interface 156, make transition passage section 122 between blade root channel section 120 and shank part 114, outwards gradually become taper.And transition passage segment length 146 makes and limit an arc-shaped interface 170 between transition passage section 122 and shank part 114.
Rotor blade 100 is known, is similar to blade 42 substantially.Correspondingly, blade 100 has platform 52, shank 54 and tenon 56.In addition, blade 100 also comprises a cooling cavity 202, and it is substantially similar to cooling cavity 102, and is defined by the internal surface 204 of blade 100.Blade cooling cavity 202 also comprises airfoil portion 116, tenon part 212 and the shank part 114 that flows and link to each other, and the cooling fluid of sending into cooling cavity tenon part 212 like this enters cooling cavity airfoil portion 116 by part 212 and 114.Cooling cavity tenon part 212 comprises a blade root channel section 220 and a transition passage section 222 that flows and link to each other.More specifically, blade root channel section 220 extends between blade end 101 and transition passage section 222, and transition passage section 222 is extended between blade root channel section 220 and shank part 114.
Blade root channel section radial inner end 138 is near cooling cavity throat 141, and blade root channel section radial outer end 140 is near transition passage section 222.Cooling cavity throat 138 blade end 101 be limited to the lower blade tang to 86 between, blade root channel section radial outer end 140 is limited to the upper blade tang between 84.
Transition passage section 222 expands outwardly to cooling cavity shank part 114 from blade root channel section 220.Therefore, the width 240 of transition passage section 222 changes between the radial inner end 242 of transition passage section 222 and radial outer end 244.Transition passage section width 240 is greater than blade root channel section width D RThe length 246 of transition passage section 222 is measured between transition passage section end 242 and 244.Because length 246 is less than transition passage length 146, so transition passage section 222 suddenly expands outwardly to shank part 114 from blade root channel section 220, so that transition passage section width 240 equals shank partial width DS at 244 places, transition passage section end.Suddenly the result of transition forms a bottom angle of cut 256 between transition passage section 222 and blade root channel section 220, forms a top angle of cut 258 between transition passage section 222 and shank part 114.And, because length 246 less than transition passage segment length 146, has formed the top angle of cut 258 between the upper blade tang is to 84.
In the manufacture process of blade 42, the anti-oxidant environment coating of aerofoil internal surface 104 coating one decks.In one embodiment, this anti-oxidant environment coating be a kind of can be from the Howmet of Michigan Whitehall, the aluminide coating that Thermatech has bought.In this exemplary embodiment, anti-oxidant environment coating is utilized the vapor aluminide deposition process and is coated on the aerofoil internal surface.Arc- shaped interface 156 and 170 and the combination of transition passage section 122 make the applied thickness of anti-oxidant environment coating greater than reaching applied thickness in the blade 100.Particularly, in blade 100, the thickness limit of known environment coating is less than 0.001 inch.Yet in blade 42, the thickness of coating can reach 0.015 inch.The thickness of this increase makes the manufacturing coating control that is used to limit the thickness that is coated to the coating on the blade 100 require to be lowered in the manufacturing of blade 42, compares with blade 100 like this, and the overall manufacture cost of blade 42 has just reduced.
In the manufacture process of cavity 102, a core (not shown) is cast in the blade 42.This core is made by a kind of liquid ceramics of injection or graphite oar in a core mould (not shown).This oar is heated to form a solid ceramic aerofoil core.This aerofoil core is suspended in the aerofoil mould, and hot wax is injected the aerofoil mould to center on ceramic aerofoil core.Hot wax solidifies, and formation one has the aerofoil (not shown) of the ceramic core that is suspended in the aerofoil.
The wax system aerofoil that then will have ceramic core immerses in the ceramic oar liquid, makes its drying.Repeat this operation for several times, so that on wax system aerofoil, form a housing.Then wax melts away and remaining has the model that is suspended on its inner core from housing, molten metal is injected in it again, metal-cured after, housing is destroyed, take out core.
In the engine operation process, cooling fluid enters blade 42 by cooling cavity blade root channel section 120.In one embodiment, cooling fluid is infeeded blade 42 from gas compressor such as gas compressor 14 (as shown in Figure 1).The cooling fluid that enters into blade tenon 56 is flowed through blade root channel section 120 and transition passage section 122 and is entered cooling cavity shank part 114.Then cooling fluid enters the cooling chamber 108 that is limited in the cooling cavity airfoil portion 116.When combustion gas strikes on the blade 42, on blade internal surface 104, produce a running temperature.Even anti-oxidant environment coating also helps reducing the oxidation on blade interior surface 104 under the situation that running temperature rises.
And at run duration, the stress that engine operation produced can act on the blade tenon 56.Compare with blade 100, the increase of the thickness of the anti-oxidant environment coating in the blade 42 helps preventing the degeneration of the material in the blade tenon 56, thereby guarantees the fatigue life of blade 42.More specifically, arc- shaped interface 156 and 170 helps the appearance of the crackle of the anti-oxidant environment coating in the limit blade tenon 56, therefore the working life that can improve blade 42.And, in running, to compare with 258 with the angle of cut 256 of blade 100, arc- shaped interface 156 and 170 also helps reduction and may lead the operation stress of tenon 56 thereby the working life that helps improving blade 42 equally.
Above-mentioned blade is not only highly reliable but also to one's profit.This blade has the local at least interior cooling cavity of tenon part that is limited to blade.The cooling cavity that is limited in the tenon has arc-shaped transition between its each several part.These arc-shaped transition are compared with rotor blade of the prior art and are helped reducing the operation stress of tenon of may leading.In addition, compare with the prior art blade, arc-shaped transition can make in the thicker anti-oxidant environment coating of internal surface coating of blade.Arc-shaped transition helps reducing the appearance than the crackle of thick coating in the blade tenon.Therefore, the geometry design of blade of the present invention and environment coating to calculate and reliably mode help working life of guaranteeing thermal fatigue life and improving aerofoil.
Though above formal description with specific embodiment the present invention, those skilled in the art can expect other substitute mode of the present invention in the scope of aim of the present invention and claim.
Component list 10 gas turbines 12 blower fan apparatus 14 high-pressure compressors 16 combustion chambers 18 high-pressure turbines, 20 low-pressure turbines, 28 entrance sides, 30 outlet sides, 40 rotor arrangements, 42 rotor blades, 44 rotor disks, 50 aerofoils, 52 platforms, 54 shanks, 56 tenons, 60 the first side walls, 62 second sidewalls, 64 guide lugs, 66 trailing edges, 68 blade roots, 70 blade tips, 74 outer surfaces, 80 tang, 82 tongue-and-grooves, 84 top tang cool off cavitys 104 inner surfaces 106 inwalls 108 cooling chambers 112 cavity tenon parts 114 cavity shank parts 116 cavity airfoil portions 120 tenon part blade root channel sections 122 tenon part transition passage section D to 86 bottom tang to 100 prior art rotor blades, 101 blades inner 102RThe blade root channel section width 132 negative pressure sidewalls 134 pressure sidewalls 136 length 138 radial inner end 140 radial outer ends 141 cooling cavity D of throatSShank partial width DTTransition passage section width 142 radial inner end 144 radial outer ends 146 length 156 arc-shaped interfaces 170 arc-shaped interfaces 202 cooling cavitys 204 inner surfaces 212 tenon parts 220 blade root channel sections 222 transition passage sections 240 transition passage section width 242 radial inner end 244 radial outer ends 246 length

Claims (20)

1. the manufacture method of the blade (42) of a gas turbine (10), wherein this blade has an aerofoil (50), a platform (52), a shank (54) and a tenon (56), platform extends between aerofoil and shank, shank extends between tenon and platform, tenon comprises that at least one is used for vanes fixed in in-engine tang (80), and this method comprises:
Limit a cooling cavity (102) in blade, it extends through aerofoil, platform, shank and tenon, wherein is limited to cavity in the tenon and partly comprises one and have the first width (D R) blade root channel part (124), and one at the blade root passage be limited to the transition portion (122) that extends between the cavity part (114) in the shank, is limited to the second width (D that the cavity part in the shank is had S) greater than first width of blade root passage; With
The anti-oxidant environment coating of coating one deck at least a portion of the internal surface (104) that limits this blade that cools off cavity.
2. the method for claim 1 is characterized in that the step that limits cooling cavity (102) also is included in qualification cooling cavity in the tenon (56), so that the blade root passage first width (D R) in root passage (124), be constant substantially, and make transition portion (122) at the root passage be limited between the cavity part (114) in the shank taperedly, make the width (D of transition portion S) be transformable in transition portion.
3. method as claimed in claim 2, it is characterized in that the step that limits cooling cavity (102) also comprises this cooling cavity of qualification, so that interface (170) formation one between tenon transition portion (122) and the shank part (54) limits the arc of this part cooling cavity.
4. method as claimed in claim 2 is characterized in that the step that applies also is included in applied thickness at least a portion of internal surface of the cooling cavity (102) in the tenon (56) greater than 0.001 inch coating at least a portion of blade (42) internal surface (104).
5. the method for claim 1 is characterized in that the step that applies also is included in coating on the internal surface of the cooling cavity (102) in the tenon (56) and is beneficial to reduce the fatigue crack that periodic duty produced in the tenon at least a portion of blade (42) internal surface (104).
6. the blade (42) of a gas turbine (10), this blade has:
One platform (52);
One shank (54) from this platform extension; With
One extends between end of this blade and shank, is used for the tenon (56) of blade installation in gas turbine, and this tenon has at least one tang (80);
One aerofoil (50), this aerofoil have a first side wall (60) and one second sidewall of stretching along the radially leaf extension between platform and the blade tip (70) (62); With
One is limited to cooling cavity (102) in the blade by tenon, shank, platform and aerofoil, this cooling cavity has a tenon part (112) that is limited in the tenon, the one shank part (114) and that is limited in shank and the platform is limited to the interior airfoil portion (116) of aerofoil, wherein the shank part becomes to flow continuous between airfoil portion and tenon part, tenon partly comprises a blade root passage (124) and a transition passage (122), and the blade root passage has one first width (D R), shank partly has the second width (D greater than blade root passage first width S), and transition passage is connected between root passage and the shank part.
7. blade as claimed in claim 6 (42) is characterized in that the described cooling cavity blade root passage first width (D R) between the suction side (132) of on the pressure side (134) of cooling off cavity (102) and described cooling cavity, measure described blade root passage first width substantial constant in described blade root passage (124).
8. blade as claimed in claim 6 (42) is characterized in that the described cooling cavity shank channel second width (D S) between the suction side (132) of on the pressure side (134) of cooling off cavity (102) and described cooling cavity, measure, the interface (170) of described transition passage (122) and shank part (114) is an arc.
9. blade as claimed in claim 8 (42) is characterized in that described cooling cavity interface (170) helps being reduced in the working stress that produces in the blade tenon (56).
10. blade as claimed in claim 6 (42) is characterized in that described tenon (56) also comprises the internal surface (104) that limits cooling cavity tenon part (112), the anti-oxidant environment coating of described tenon internal surface coating one deck.
11. blade as claimed in claim 6 (42) is characterized in that described tenon (56) also comprises the internal surface (104) that limits cooling cavity tenon part (112), it is 0.001 inch anti-oxidant environment coating that described tenon internal surface applies a layer thickness.
12. blade as claimed in claim 6 (42) is characterized in that the structure of described cooling cavity (102) helps reducing the appearance of tenon low cycle fatigue crackle.
13. gas turbine (10) with a plurality of blades (42), each blade has an aerofoil (50), one shank (54) and a platform that between aerofoil and shank, extends (52), each blade also has a cooling cavity (102) and and has at least one tang (80) and be used for each vanes fixed in in-engine tenon (56), shank extends between platform and tenon, the cooling cavity is limited at aerofoil, platform, in shank and the tenon, described cooling cavity has tenon part (112), one shank part (114) with its airfoil portion (116) that flows and link to each other, it is the first width (D that described cooling cavity tenon partly has its width R) a blade root passage (124) and a transition passage (122), described cooling cavity shank partly has one second width (D S), it is greater than first width of blade root passage, and described cooling cavity transition passage is tapered between root passage and shank part.
14. gas turbine as claimed in claim 13 (10) is characterized in that the described cooling cavity blade root passage first width (D R) between on the pressure side (134) of cooling off cavity (102) and suction side (132), measure the described cooling cavity shank part second width (D S) between described cooling cavity pressure side and suction side, measure described passage first width substantially constant in described passage.
15. gas turbine as claimed in claim 14 (10) is characterized in that the interface (170) between described cooling cavity transition passage (122) and the cooling cavity shank part (114) has a radius.
16. gas turbine as claimed in claim 14 (10) is characterized in that described cooling cavity (102) coating has certain thickness anti-oxidant environment coating to reduce the low cycle fatigue of each blade (42).
17. gas turbine as claimed in claim 14 (10), at least a portion applied thickness that it is characterized in that described cooling cavity (102) is greater than 0.001 inch anti-oxidant environment coating.
18. gas turbine as claimed in claim 14 (10), at least a portion applied thickness that it is characterized in that described cooling cavity tenon part (112) is greater than 0.001 inch anti-oxidant environment coating.
19. gas turbine as claimed in claim 14 (10) is characterized in that each described cooling cavity (102) helps reducing the working stress that produces in each described blade tenon (56).
20. gas turbine as claimed in claim 14 (10) is characterized in that the structure of described cooling cavity (102) helps reducing the appearance of tenon low cycle fatigue crackle.
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JP4458772B2 (en) 2010-04-28
CN100572757C (en) 2009-12-23
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EP1365108A2 (en) 2003-11-26
US20030219338A1 (en) 2003-11-27
EP1365108A3 (en) 2004-10-06

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